Siemens Energy, Inc.

Siemens Energy Global Gmbh & Co. Kg

Siemens Aktiengesellschaft

Jose L Rodriguez

Ali Akturk

Gary Brian Merrill

Matthew D Montgomery

Matthew J Golsen

Ian T Doran

Jay Alan Morrison

David James Wiebe

Alexander Ralph Beeck

Richard Colin Charron

David Allen Little

John A Orosa

Jiping Zhang

Abdullatif M Chehab

Megan Schaenzer

Bernhard W Küsters

David L Wasdell

Marco Claudio Pio Brunelli

Lara Richerson

David Monk

James O'Brien

John T Harrington

Joseph L Evins

Timothy Troyer

John Harrington

Krishan Mohan

Taylor Hynds

Louis Charles

Jesus Velez-Quinones

Fuad Ismayilov

Ardit Bitincka

Stephen Wright

Gary B Merrill

A waste heat recovery systemis provided to (i) capture waste heat, such as from or within an exhaust pipeof a tractor-trailer, (ii) extract a thermal heat value from the waste heat, (iii) convert the thermal heat value into electricity, and (iv) route the electricity to an electricity sink, such as a tractor-trailerTransport Refrigeration Unit (TRU).

Waste heat recovery system

A ring segment assembly includes a ring segment including an impingement pocket having an impingement surface, a plurality of pins extending from the impingement surface, and an impingement plate spaced a non-zero distance from the impingement surface. The plurality of pins are arranged to define a plurality of pinless impingement areas. The impingement plate has a plurality of bumps and a plurality of valleys. The impingement plate defines a plurality of impingement holes. Each impingement hole of the plurality of impingement holes is formed in one of the valleys of the plurality of valleys and positioned opposite one of the plurality of pinless impingement areas.

Ring segment assembly in gas turbine engine

An investment casting core () incorporates an alignment guide () extending through a body () of the core. The alignment guide () defines a coolant flow path () in a later-cast metal component () extending from a coolant outlet opening () in an impingement structure () to an impingement target area () of a cooling feature () formed on an impingement cooled surface () of the component (). Methods of making the core () and using the core () in lost wax investment casting processes are also described.

Investment casting core with cooling feature alignment guide and related methods

A turbine blade in a gas turbine engine includes an airfoil extending in a radial direction. The airfoil has an outer wall delimiting an airfoil interior. The outer wall includes a pressure sidewall and a suction sidewall joined at a leading edge and a trailing edge in a longitudinal direction. A turbulator is disposed in the airfoil interior. The turbulator includes a first row having at least two turbulator ribs spaced apart in the longitudinal direction. The turbulator includes a second row extending in the radial direction from the first row and having at least two turbulator ribs spaced apart in the longitudinal direction.

Turbine blade in gas turbine engine

A molding tool () for manufacturing cooling features in a ceramic core for a casting process includes a first mold portion () comprising a crossover hole forming feature (); a second mold portion () comprising an impingement jet receiving forming feature () for forming an impingement jet receiving feature having a desired aiming point (); and a sacrificial alignment member () for extending at least partially into the crossover hole forming feature () at least partially into the aiming point () of the impingement jet receiving forming feature () for substantially aligning a central axis () of the crossover hole forming feature () with the aiming point () to maintain a crossover hole and aiming point alignment during casting to ensure that the jet is directed at the aiming point () in a resultant cast part.

Manufacturing aligned cooling features in a core for casting

A turbine blade tip includes a tip cap disposed over a blade airfoil and having a pressure side edge and a suction side edge. A notch is formed by a radially inward step adjacent to the suction side edge of the tip cap. The notch is defined by a radially extending step wall and a radially outward facing land. The step wall extends radially inward from the suction side edge of the tip cap to the land, whereby the land is positioned radially inward in relation to a radially outer surface of the tip cap. The notch extends along at least a portion of the suction sidewall in a direction from the leading edge to the trailing edge. In a further aspect, a method is provided for servicing a blade that includes machining a suction side notch as described above.

Turbine blade and corresponding servicing method

An airfoil for a turbine engine includes an array of pins positioned in an internal cavity of the airfoil, such that cooling channels are defined in the interspaces between adjacent pins. Each pin extends lengthwise from a first airfoil wall to a second airfoil wall and is connected thereto at a first intersection and at a second intersection respectively. The pin has a first cross-sectional shape at a respective intersection and a second cross-sectional shape at an intermediate plane located between the first and second intersections. The first cross-sectional shape includes a closed shape defined by relatively sharp corners and the second cross-sectional shape includes a closed shape defined by relatively rounded corners. A cross-sectional area of the pin at the intermediate plane is greater than a cross-sectional area of the pin at the respective intersection.

Airfoil for a turbine engine incorporating pins

An impingement cooling system for a gas turbine engine includes an initial impingement surface () with a centrally located opening (). A plurality of channels () and plurality of sub-channels () extends radially outward from the opening () and are formed by a plurality of fixtures () and plurality of sub-fixtures () that each separates each adjacent channel () and sub-channel () respectively. The plurality of fixtures () and plurality of sub-fixtures () each have a rounded upstream end () in a plane parallel relative to the initial impingement surface (). The plurality of fixtures () and the plurality of sub-fixtures () each have a concave shape along a middle portion () of the fixture () and sub-fixture () along an axis perpendicular to the initial impingement surface (). The plurality of channels () is divided into the plurality of sub-channels () extending radially outward of an inlet of each channel () from a stagnation point () created in the channel at an upstream end () of the sub-fixture ().

Impingement cooling features for gas turbines

An arrangement to minimize vibrations in a gas turbine exhaust diffuser is provided. The arrangement includes a projection coupled to an inner cylindrical surface or the outer cylindrical surface of a fluid flow path of the gas turbine exhaust diffuser. The projection minimizes pressure oscillations in the gas turbine exhaust diffuser such that an unsteadiness of the fluid flow surrounding the second tangential strut is reduced. A method to minimize pressure oscillations in a gas turbine diffuser is also provided.

Gas turbine diffuser outer diameter and inner diameter wall strips for turbine exhaust manifold pressure oscillation reduction

A midframe portion () of a gas turbine engine () is provided, including a compressor section () with a last stage blade (). The compressor section () is configured to introduce a radial velocity component into an air flow () such that the air flow is discharged from the compressor section () at a mixed direction based on a combined longitudinal velocity component and radial velocity component. The midframe portion () further includes a manifold () to directly couple the air flow from an outlet of the compressor section () to an inlet of a respective combustor head () of the gas turbine engine ().

Mid-section of a can-annular gas turbine engine to introduce a radial velocity component into an air flow discharged from a compressor of the mid-section

A combustor assembly () including guide vanes () located between an inner cylinder () and a flow sleeve (). Each guide vane () includes a circumferentially angled flow directing portion () adjacent to a leading edge (). The leading edge () of at least one guide vane () can be located radially inward along the longitudinal axis () relative to the leading edge () of at least one other of the guide vanes (). The length of the guide vanes () may vary, and the circumferential spacing between a first pair of the guide vanes () can be different from a spacing between a second pair of the guide vanes ().

Combustor arrangement including flow control vanes

An arrangement to control vibrations in a gas turbine exhaust diffuser is provided. The arrangement includes a protrusion coupled to a turbine exhaust cylinder strut for controlling shock induced oscillations in a gas turbine diffuser. The controlled shock induced oscillations minimize pressure fluctuations in the gas turbine exhaust diffuser such that an unsteadiness of the fluid flow surrounding the turbine exhaust cylinder strut is reduced. A method to fluid flow induced vibrations in a gas turbine diffuser is also provided.

Turbine exhaust cylinder strut strip for shock induced oscillation control

A turbine airfoil usable in a turbine engine and having at least one cooling system with an efficient trip strip is disclosed At least a portion of the cooling system may include one or more cooling channels having one or more trip strips protruding from an inner surface forming the cooling channel. The trip strip may have improved operating characteristics including enhanced heat transfer capabilities and a substantial reduction in pressure drop typically associated with conventional trip strips In at least one embodiment, the trip strip may have a cross-sectional area with a first section of an upstream surface of the trip strip being positioned nonparallel and nonorthogonal to a surface forming the cooling system channel extending upstream from the at least one trip strip and a concave shaped downstream surface of the at least one trip strip that enables separated flow to reattach to the cooling fluid flow.

Turbine airfoil with an internal cooling system having trip strips with reduced pressure drop

A gas turbine engine can-annular combustion arrangement (), including: an axial compressor () operable to rotate in a rotation direction (); a diffuser () configured to receive compressed air () from the axial compressor; a plenum () configured to receive the compressed air from the diffuser; a plurality of combustor cans () each having a combustor inlet () in fluid communication with the plenum, wherein each combustor can is tangentially oriented so that a respective combustor inlet is circumferentially offset from a respective combustor outlet in a direction opposite the rotation direction; and an airflow guiding arrangement () configured to impart circumferential motion to the compressed air in the plenum in the direction opposite the rotation direction.

Swirling midframe flow for gas turbine engine having advanced transitions

A turbine engine temperature control system configured to limit thermal gradients from being created within an outer casing surrounding a turbine airfoil assembly during shutdown of a gas turbine engine and for preheating an engine during a cold startup is disclosed. By reducing thermal gradients caused by hot air buoyancy within the mid-region cavities in the outer casing, arched and sway-back bending of the outer casing is prevented, thereby reducing the likelihood of blade tip rub, and potential blade damage, during a warm restart. The turbine engine temperature control system may also be used for cold startup conditions to heat engine components such that gaps between turbine airfoil tips and adjacent blade rings can be made larger from thermal expansion, thereby reducing the risk of damage. The turbine engine temperature control system may operate during turning gear system operation after shutdown of the gas turbine engine or during a cold startup.

Growing community of inventors

Lake Mary, FL, United States of America

Jose L Rodriguez